1. Technical Field
The present invention relates to gas turbine engines and, more particularly, to a brush seal for minimizing air leakage therein.
2. Background Art
A typical gas turbine engine includes a low compressor, a high compressor, a combustor, a high turbine, and a low turbine sequentially situated about a longitudinal axis. The sections of the gas turbine engine are enclosed in an engine case with air flowing axially through the engine along an annular air path. As is well known in the art, the air enters the low compressor at a substantially ambient temperature and pressure and is compressed by the low and high compressor sections, respectively, to elevate its temperature and pressure. The compressed air is then mixed with fuel, ignited, and burned in the combustor. The hot products of combustion emerging from the combustor are expanded in the high and low turbines, respectively, thereby rotating the turbine and driving the compressor.
The high turbine section of the gas turbine engine is subjected to extremely harsh environment, characterized by very high temperatures and pressures. The components of the high pressure turbine must be cooled to prevent these components from burning in a very short period of time. The cooler air is typically bled from the high pressure compressor and routed to the high pressure turbine. Although the bled cooling air is necessary to cool certain engine parts, loss of the cooling air from the compressor is highly undesirable. The cooling air, diverted from the compressor, is no longer available to produce thrust and thus negatively effects the efficiency of the gas turbine engine. Therefore, the cooling air is carefully routed from the high pressure compressor to the high pressure turbine to ensure that this valuable cooling air is not wasted.
The cooling air is routed through the engine into cooling air compartments that are substantially adjacent to the high pressure turbine air path. The cooling air from the cooling air compartments is carefully metered out to cool various components of the turbine section of the gas turbine engine. However, a barrier is required to prevent the cooling air from escaping from the cooling air compartment into the turbine air path, other than through designated bleeding holes, since the cooling air is at a greater pressure than the air in the turbine air path.
The conventional barriers are annular metal seals which are not completely effective for a number of reasons. First, certain sealing surface which abut these seals are not uniform. For example, a plurality of blade outer air seals forms a segmented ring around the tips of rotating blades. Even slight size differences among the blade outer air seals results in gaps between some of the blade outer air seals and the barrier seals, thereby allowing some of the cooling air to escape therethrough. A second difficulty with effectively sealing the cooling air compartments located in the high pressure turbine of the gas turbine engine is that the high pressure turbine experiences great temperature fluctuations. As a result of these large temperature gradients, the turbine components are subjected to significant thermal expansion. When components of the high pressure turbine expand, the conventional seals cannot adjust and seal the gap between expanded components and the seal, therefore, allowing some of the cooling air to escape through these gaps. Therefore, there is a great need to effectively seal cooling air compartments in the high pressure turbine.